The present invention generally relates to orbit inclination and altitude control for spacecraft and, more particularly, to a satellite orbit inclination control system with redundancy to enhance reliability.
It is common for satellites in orbit to use two north and two south pointing electric plasma thrusters for control of the inclination of the satellite""s orbit, and for station keeping and altitude control. For example, FIG. 1 shows satellite 100, which may be a geosynchronous satellite, in two different positions in its flight orbit 102 about the earth 104. A geosynchronous orbit is of particular importance for many satellites, including communications and navigation satellites. FIG. 1 shows geosynchronous orbit 106 in the plane 108 of the equator 110 of the earth 104. Because equator 110 of the earth 104 lies in plane 108, plane 108 may also be referred to as the equatorial plane, or the equatorial orbit plane of geosynchronous orbit 106. As seen in FIG. 1, flight orbit 102 of satellite 100 lies in a plane which is inclined to equatorial orbit plane 108, as indicated in FIG. 1 by angle of inclination xcex8. In order to maintain satellite 100 in geosynchronous orbit, it is desired to provide inclination control by actuating north and south pointing thrusters at nodes 112 and 114, where the plane of flight orbit 102 and the equatorial orbit plane 108 of geosynchronous orbit 106 intersect. For example, by actuating a north pointing thruster at node 112, a correction in the direction indicated by arrow 116 may be provided, and by actuating a south pointing thruster at node 114, a correction in the direction indicated by arrow 118 may be provided, in order to correct the inclination of flight orbit 102 back to a geosynchronous orbit 106 in equatorial orbit plane 108. In other words, north and south pointing thrusters are actuated at nodes 112 and 114 to reduce angle of inclination xcex8 to approximately 0.0 degrees.
The preferred mounting configuration for north and south pointing thrusters on satellite 100 is generally anti-nadir side 120 of satellite 100, i.e., the side of satellite 100 facing away from, or furthest from, the earth 104. Anti-nadir side 120 may be the preferred configuration for mounting thrusters due to various engineering constraints and for other technical reasons. For example, mounting thrusters on anti-nadir side 120 of satellite 100 may minimize interference of the thrusters with radio antennas and other communications devices requiring an unobstructed path to the earth 104, in the case where satellite 100 is a communications satellite. Because the force, or thrust, applied by north and south pointing thrusters to satellite 100 must be directed through the center of mass of satellite 100 to prevent torqueing satellite 100, there is a radial component to the thrust which acts to change the eccentricity of flight orbit 102. By actuating first one thruster at node 112 and then actuating an oppositely pointing thruster at node 114 on the opposite side of flight orbit 102, for example, first a north pointing and then a south pointing thruster, the radial component of the thrust may be effectively cancelled so as not to change the eccentricity of flight orbit 102, but only the angle of inclination xcex8 of flight orbit 102. Thus, an operational system for inclination and altitude control of satellite 100 requires, at a minimum, a north and a south pointing thruster.
To ensure mission success, thrusters and their control electronics are currently provided with backup units. For example, where a minimum of two thrusters, i.e., a north pointing thruster and a south pointing thruster, is required, four thrusters may be installed to guarantee access to two thrusters in case of a single thruster failure. The approach of providing more thrusters, or in general more components of any type, than the minimum required in order to enhance reliability, guarantee access to the minimal number of components required in case of single component failure, and to ensure mission success is known as redundant design, or more briefly, redundancy.
Electric propulsion, more specifically plasma type electric propulsion, has been introduced for satellite control as a replacement for chemical propulsion primarily because of the improved specific impulse, i.e. the change in momentum produced using a unit mass of propellant, of plasma type electric propulsion over chemical propulsion. The specific impulse of electrical plasma thrusters is approximately an order of magnitude, or 10 times, greater than the specific impulse of chemical thrusters. Chemical propulsion thruster systems use many thrusters (typically 12) with optimal thrust orientations to reduce propellant use. A number of factors, such as component mass, mounting space and exhaust plume size, for example, do not allow for a simple one-for-one exchange of electrical plasma thrusters for chemical thrusters. In the example that follows, a chemical propulsion thruster system using 12 thrusters is compared to an electrical plasma thruster system using a reduced number of electrical plasma thrusters, i.e., 4 electrical plasma thrusters.
FIG. 2 shows an example of an electrical plasma thruster system for satellite 100 using electrical plasma thrusters 122, 124, 126, and 128 as typically currently mounted at the anti-nadir side 120 of satellite 100. Two-axis gimbals mechanisms 130 and 132 are used to align the thrust vectors of electrical plasma thrusters 122, 124, 126, and 128 to the center of mass of satellite 100. This arrangement for an electrical plasma thruster system may reduce the overall weight of satellite 100, including weight of propellant required for some particular desired life span of satellite 100 and component weight of the thruster system, compared to a chemical propulsion thruster system. The cost of the electrical plasma thruster system, however, may be greater.
For example, in a typical satellite weighing approximately 3,000 kilograms (Kg) with a desired lifespan of approximately 15 years, approximately 1,000 Kg propellant may be needed for chemical thrusters compared to only 100 Kg of propellant for electrical plasma thrusters mounted in the same arrangement as the chemical thrusters. The propellant savings is primarily due to the order of magnitude advantage in specific impulse for the electrical plasma thrusters over chemical thrusters. The two-axis gimbals mounting arrangement using only 4 electrical plasma thrusters instead of 12 is less fuel efficient, however, requiring thruster actuations on opposite sides of the flight orbit as described above, so that 150 Kg of propellant may be needed for the 4 electrical plasma thrusters.
Additional mass inefficiency may be incurred when system redundancy is achieved by duplication of all components, rather than duplication of the functional requirements. Each electrical plasma thruster with its electronic controller and gimbals mechanism typically weighs 40 Kg compared to approximately 0.5 Kg for a chemical thruster. Thus, 12 chemical thrusters may be expected to weigh approximately 6 Kg, a negligible amount, compared to 480 Kg for 12 electrical plasma thrusters. By reducing the number of electrical plasma thrusters to four, the weight penalty for the electrical plasma thrusters is reduced to approximately 160 Kg. Thus, the total propellant and component weight for a chemical propulsion system may be expected to be approximately 1,000 Kg compared to a total weight of approximately 310 Kg for an electrical plasma thruster system using only 4 electrical plasma thrusters.
Typical cost for a chemical thruster system with 12 thrusters may be expected not to exceed approximately $1.5 million whereas typical cost for an electrical plasma thruster system with 4 thrusters may be expected not to exceed approximately $4.0 million. A simple one-for-one exchange of electrical plasma thrusters for chemical thrusters, then, becomes prohibitively expensive. In summary, a weight savings of approximately 700 Kg may be achieved, but at a cost penalty of approximately $2.5 million, by replacing a chemical propulsion thruster system with an electrical plasma thruster system having a reduced number of thrusters.
As can be seen, there is a need for an electrical plasma thruster system for inclination control, station keeping, and altitude control of satellite and spacecraft orbits, which reduces the weight and expense of the thruster system over the conventional two-axis gimbals mounting arrangements for electrical plasma thruster systems. There is also a need for an electrical plasma thruster system for inclination control, station keeping, and altitude control of satellite and spacecraft orbits which assures and improves reliability of the thruster system by providing the required redundancy.
The present invention provides an electrical plasma thruster system for inclination control, station keeping, and altitude control of satellite and spacecraft orbits, which reduces the weight and expense of the thruster system over the conventional mounting arrangements of four thrusters for electrical plasma thruster systems. The present invention also provides an electrical plasma thruster system for inclination control, station keeping, and altitude control of satellite and spacecraft orbits which assures and improves reliability of the thruster system by providing required redundancy, as well as by substituting more reliable components for less reliable ones.
In one aspect of the present invention, a system includes a thruster, a thruster mounting boom, a gimbals mechanism connecting the thruster to the thruster mounting boom, and a pivot mechanism connected to the thruster mounting boom, where the pivot mechanism attaches the thruster mounting boom to a spacecraft.
In another aspect of the present invention, a redundant system includes a first electrical plasma thruster and a second redundant electrical plasma thruster, a thruster mounting boom, a two-axis gimbals mechanism connecting the first thruster to the thruster mounting boom, a pivot mechanism connected to the thruster mounting boom, where the pivot mechanism attaches the thruster mounting boom to the spacecraft.
In still another aspect of the present invention, a redundant system includes a first electrical plasma thruster; a first thruster mounting boom; a first two-axis gimbals mechanism connecting the first thruster to the first thruster mounting boom, where the redundant system includes a first actuator and a first control unit for controlling the first two-axis gimbals mechanism to position and point the first electrical plasma thruster; and a first pivot mechanism connected to the first thruster, mounting boom, where the first pivot mechanism attaches the first thruster mounting boom to the anti-nadir side of the spacecraft, and the redundant system includes a second actuator and a second control unit for pivoting and positioning the first thruster mounting boom. The redundant system also includes a second electrical plasma thruster; a second thruster mounting boom; a second two-axis gimbals mechanism connecting the second electrical plasma thruster to the second thruster mounting boom where the redundant system includes a third actuator and a third control unit for controlling the second two-axis gimbals mechanism to position and point the second electrical plasma thruster; and a second pivot mechanism connected to the second thruster mounting boom, where the second pivot mechanism attaches the second thruster mounting boom to the anti-nadir side of the spacecraft, and the redundant system includes a fourth actuator and a fourth control unit for pivoting and positioning the second thruster mounting boom.
In a further aspect of the present invention, a method includes steps of providing a first thruster mounted on a pivoting thruster mounting boom attached to a spacecraft; using the first thruster at a first position, for example, a north pointing position, to provide orbit control; repositioning the first thruster to a second position, for example, a south pointing position; and using the first thruster at the second position to provide orbit control so as to provide redundancy for a failed second thruster by repositioning the first thruster.
These and other features, aspects and advantages of the present invention will become better understood with reference to the following drawings, description and claims.